A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine conventionally includes, for example, five major sections: a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. The fan section is typically positioned at the inlet section of the engine and includes a fan that induces air from the surrounding environment into the engine and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum and out the exhaust section.
The compressor section raises the pressure of the air it receives from the fan section, and the resulting compressed air then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel into a combustion chamber formed between inner and outer liners. The fuel and air mixture is ignited to form combustion gases, which drive rotors in the turbine section for power extraction. The gases then exit the engine at the exhaust section.
In a typical configuration, the turbine section includes rows of stator vanes and rotor blades disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft.
In most gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components in order to prevent overheating and potential mechanical failures attributable thereto. As such, most turbine components, particularly the stator vane and rotor blade assemblies may benefit from temperature management in view of the high temperature environment of the mainstream hot gas flow path. Accordingly, in many turbine sections, the volumetric space disposed radially inwardly or internally from the hot gas flow path includes an internal cavity through which a cooling air flow is provided. The cooling of the internal engine cavity attempts to maintain the temperatures of the rotor disks and other internal engine components that are suitable for their material and stress level.
However, in many conventional engines, relatively high levels of cooling air flows have been used to obtain satisfactory temperature control of the components within the internal engine cavity. In addition, the demand for cooling flow may be impacted by an irregular and unpredictable ingestion of mainstream hot gases from the hot gas flow path into the internal engine cavity. Various attempts to prevent hot gas ingestion between adjacent stator vanes and rotor blades have primarily involved the use of overlapping lip-type structures in close running clearance, often referred to as flow discouragers, but these structures have not been as effective as desired. Moreover, it is generally desirable to employ mechanisms to minimize this cooling air since air from the compressor used for cooling is not available for combustion. Additionally, temperature control of the flow discouragers should also be considered. If the flow discouragers are exposed to undesirably high temperatures, they may deform, which may impact their primary functions.
Accordingly, it is desirable to provide an improved gas turbine engine assembly that maintains proper temperature control. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.